The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in several turbine stages which power the compressor and produce useful work such as powering an upstream fan in a turbofan aircraft engine application.
Each turbine stage includes a stationary turbine nozzle which directs the combustion gases through a corresponding row of turbine rotor blades extending radially outwardly from a supporting rotor disk.
The stator vanes and rotor blades include hollow airfoils with internal cooling circuits therein which use air bled from the compressor for cooling thereof during operation. Each stage of vanes and blades is configured differently for maximizing energy extraction from the combustion gases as they flow downstream through the turbine during operation. Each vane and blade also includes different cooling configurations specifically tailored to different heat loads from the combustion gases as they flow downstream over the pressure and suction sides thereof.
The typical turbine blade includes an airfoil having a radially outer tip spaced closely adjacent to a surrounding turbine shroud forming a small clearance or gap therebetween. During operation, the combustion gases flow over the pressure and suction sides of the turbine blade, and a small portion of the combustion gases leaks past the blade tip through the small tip clearance.
The blade tip is particularly difficult to cool since it is exposed to the hot combustion gases on both the pressure and suction sides of the airfoil between the leading and trailing edges, as well as over the tip itself in the tip-shroud gap.
Since blade tips are subject to occasional tip rubs with the surrounding turbine shroud, the tips typically include short height squealer rib extensions of the pressure and suction sidewalls extending radially outwardly from a tip floor that defines an outwardly open tip cavity. The tip floor defines the outer boundary for the internal cooling circuits of the airfoil, and positions the squealer ribs externally thereof which further increases the difficulty of tip cooling during operation.
Accordingly, turbine blade tips are subject to oxidation over extended use in the engine, and may require corresponding repair during maintenance outages of the engine. Oxidation damage to the blade tip commonly limits the useful life of the blade during operation, as well as decreases turbine efficiency as the blade tips oxidize and wear during operation and correspondingly increase the blade tip clearance.
Two significant improvements in blade tip cooling are found in U.S. Pat. Nos. 5,261,789 and 6,672,829, both assigned to the present assignee. In the earlier patent, a tip shelf is introduced along the pressure side of the turbine blade to offset inwardly the pressure side squealer rib for enhancing tip cooling by shielding the tip with a cooling air film for protection against the radial migration of combustion gases over the tip during operation.
The later patent discloses an inclined pressure side squealer rib at the tip shelf for protecting the cooling air film created thereat for enhancing tip performance.
However, since the blade tips are subject to occasional tip rubbing with the surrounding turbine shroud, the squealer ribs are subject to abrasion wear which affects both efficiency of blade tip performance, and effectiveness of the tip shelf cooling.
Accordingly, it is desired to provide a turbine blade with enhanced tip cooling notwithstanding tip rubs.